The present invention relates to solar arrays for satellites, and more particularly to rigid panel solar arrays designed to cause minimal disturbance to their host spacecraft as the spacecraft passes through the transition zones between lightness and darkness.
Spacecraft today use a number of various types of solar panels and solar arrays, including rigid panel solar arrays and flexible panel solar arrays. Rigid arrays have a number of benefits over flexible arrays. For example, for a given solar cell total surface area, rigid arrays have lower solar torque, lower aero torque, and lower gravity gradient pitch torque. Rigid arrays also have a higher flexible mode frequency and are less expensive to design and produce than flexible arrays.
One of the drawbacks of rigid panel solar arrays consists of the solar thermal snap effects which are caused when the spacecraft passes through the transition zones between lightness and darkness (a/k/a the terminator). Solar snap, or thermal snap, can result from motion/acceleration of the solar array center-of-mass when the temperature changes rapidly as the spacecraft enters or exits an eclipse. Since the array is two-dimensional in nature, the motion of the center-of-mass can result from thermal distortions (bending) in either its longitudinal or transverse directions. Some attempts have been made to minimize the thermal snap by using composite face sheet materials having low coefficients of thermal expansion (CTE). Merely using face sheet materials having low coefficients of thermal expansion, however, does not account for the effect of the solar cells and the central core materials on thermal deformation. Also, if designed to provide minimal coefficient of thermal expansion in one direction, the composite face sheet material typically will have a significant coefficient of thermal expansion in the orthogonal direction.
It is an object of the present invention to provide a rigid panel solar array for a spacecraft which is an improvement over known rigid panel solar arrays. It is also an object of the present invention to provide an improved solar array which causes minimal disturbance to its host spacecraft as the spacecraft passes through the terminator (sunrise and sunset).
It is another object of the present invention to provide a solar array which has a lower cost and is more reliable than systems today using more complex flexible array technology.
It is still another object of the present invention to provide a rigid panel solar array which reduces solar thermal snap effects when operating in a low earth orbit (LEO) environment. It is a still further object of the present invention to provide a rigid panel solar array which has minimal bending in the longitudinal direction and only moderate bending in the transverse direction, with the disturbance effect of the moderate transverse direction bending being rendered negligible through the appropriate location and design of the attachment points (hinges) such that the distortion does not result in a shift in the panel's center of mass (CM).
These and other objects, features and benefits of the present invention will become apparent from the following description of the invention when viewed in accordance with the accompanying drawings and appended claims.